Orientation feature for swirler tube

ABSTRACT

A cooling structure for a gas turbine engine comprises a gas turbine engine structure defining a cooling cavity. A cooling component is configured to direct cooling flow in a desired direction into the cooling cavity. A bracket supports the cooling component and has an attachment interface to fix the bracket to the gas turbine engine structure. A first orientation feature associated with the bracket. A second orientation feature is associated with the gas turbine engine structure. The first and second orientation features cooperate with each other to ensure that the cooling component is only installed in one orientation relative to the gas turbine engine structure. A gas turbine engine and a method of installing a cooling structure are also disclosed.

BACKGROUND

The present disclosure relates generally to a gas turbine engine, and inparticular to a swirler tube in a gas turbine engine.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section.

In pursuit of more efficient gas turbine engines, higher turbine inlettemperatures have been relied upon to boost overall engine performance.This results in increased gas path temperatures that may adverselyaffect component life. To address this issue, dedicated cooling air isextracted from a compressor section and is used to cool the gas pathcomponents in the turbine. For example, cooling air can be directed intoa rotor cavity for cooling purposes. It is important to ensure that thecooling air is directed into the rotor cavity in a manner to provide themost effective cooling.

SUMMARY

In a featured embodiment, a cooling structure for a gas turbine enginecomprises a gas turbine engine structure defining a cooling cavity. Acooling component is configured to direct cooling flow in a desireddirection into the cooling cavity. A bracket supports the coolingcomponent and has an attachment interface to fix the bracket to the gasturbine engine structure. A first orientation feature associated withthe bracket. A second orientation feature is associated with the gasturbine engine structure. The first and second orientation featurescooperate with each other to ensure that the cooling component is onlyinstalled in one orientation relative to the gas turbine enginestructure.

In another embodiment according to the previous embodiment, one of thefirst and second orientation features comprises a male feature and theother of the first and second orientation features comprises a femalefeature that is received within the male feature.

In another embodiment according to any of the previous embodiments, themale feature comprises a tab and the female feature comprises a slot.

In another embodiment according to any of the previous embodiments, themale feature comprises a pin and the female feature comprises anopening.

In another embodiment according to any of the previous embodiments, thefirst orientation feature comprises at least one tab and the secondorientation feature comprises a rib. The rib interferes with the tab toprevent mounting the bracket in an incorrect orientation relative to thegas turbine engine structure.

In another embodiment according to any of the previous embodiments, thebracket comprises a forward bracket positioned on one side of the gasturbine engine structure and an aft bracket is positioned on an oppositeside of the gas turbine engine structure. The cooling component isassociated with the aft bracket. The forward and aft brackets eachinclude the first orientation feature.

In another embodiment according to any of the previous embodiments, thegas turbine engine structure includes an opening to define a portion ofa cooling flow path, and further includes an inlet tube associated withthe forward bracket to direct cooling flow through the opening and intothe cooling component.

In another embodiment according to any of the previous embodiments, thecooling component comprises a swirler tube that redirects cooling flowfrom a first direction defined by the inlet tube to a second directionthat is transverse to the first direction.

In another embodiment according to any of the previous embodiments, theattachment interface includes at least one fastener that secures boththe forward and aft brackets to the gas turbine engine structure.

In another embodiment according to any of the previous embodiments, thegas turbine engine structure comprises a mid-turbine frame structure.

In another embodiment according to any of the previous embodiments, thecooling component comprises a swirler tube.

In another featured embodiment, a gas turbine engine comprises amid-turbine frame located axially between a first turbine and a secondturbine. The mid-turbine frame includes an opening to define a portionof a cooling flow path into a rotor cavity. A cooling component isconfigured to direct cooling flow in a desired direction into the rotorcavity. A bracket assembly supports the cooling component and has anattachment interface to fix the bracket assembly to the mid-turbineframe. A first orientation feature is associated with the bracketassembly. A second orientation feature is associated with themid-turbine frame. The first and second orientation features cooperatewith each other to ensure that the cooling component is only installedin one orientation relative to the mid-turbine frame.

In another embodiment according to the previous embodiment, the openingin the mid-turbine frame defines a center axis. The cooling componentcomprises a swirler tube that directs cooling flow in a direction thatis non-parallel with the center axis.

In another embodiment according to any of the previous embodiments, thebracket assembly comprises a forward bracket positioned at the openingon one side of the mid-turbine frame and an aft bracket positioned atthe opening on an opposite side of the mid-turbine frame. The swirlertube is associated with the aft bracket.

In another embodiment according to any of the previous embodiments, thefirst orientation feature comprises at least one tab and the secondorientation feature comprises a rib. The rib interferes with the tab toprevent mounting the bracket in an incorrect orientation relative to themid-turbine frame.

In another embodiment according to any of the previous embodiments, thefirst orientation feature is formed on both the forward and aftbrackets. The second orientation feature is formed on each side of themid-turbine frame such that each of the forward and aft brackets canonly be mounted on the mid-turbine frame in one mounting orientation.

In another featured embodiment, a method of installing a coolingstructure in a gas turbine engine includes providing a gas turbineengine structure that defines a cooling cavity and a cooling componentthat is configured to direct cooling flow in a desired direction intothe cooling cavity. A bracket configured to support the coolingcomponent and to attach the bracket to the gas turbine engine structureis provided. A first orientation feature is formed on the bracket. Asecond orientation feature is formed on the gas turbine enginestructure. The first and second orientation features are associated witheach other to ensure that the cooling component is only installed in oneorientation relative to the gas turbine engine structure.

In another embodiment according to the previous embodiment, one of thefirst and second orientation features comprises a male feature and theother of the first and second orientation features comprises a femalefeature, and includes inserting the female feature into the male featureto achieve the one orientation.

In another embodiment according to any of the previous embodiments, oneof the first and second orientation features interferes with the otherof the first and second orientation features to prevent improperinstallation.

In another embodiment according to any of the previous embodiments, thegas turbine engine structure comprises a mid-turbine frame and thecooling component comprise a swirler tube.

The foregoing features and elements may be combined in any combinationwithout exclusivity, unless expressly indicated otherwise.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a schematic perspective view of an example mid-turbine framein the gas turbine engine.

FIG. 3A is an exploded view of a mounting assembly for a swirler tube.

FIG. 3B is a magnified view of a forward bracket of the mountingassembly of FIG. 3A.

FIG. 4 is a magnified view of an aft bracket of the mounting assembly ofFIG. 3A.

FIG. 5 is schematic view of an alternate orientation feature.

FIG. 6 is a schematic view of another alternate orientation feature.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft(10,668 meters), with the engine at its best fuel consumption—also knownas “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is theindustry standard parameter of lbm of fuel being burned divided by lbfof thrust the engine produces at that minimum point. “Low fan pressureratio” is the pressure ratio across the fan blade alone, without a FanExit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosedherein according to one non-limiting embodiment is less than about 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7 ° R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second (350.5 meters/second).

The example gas turbine engine includes fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, fan section 22 includes less than about 20 fanblades. Moreover, in one disclosed embodiment low pressure turbine 46includes no more than about 6 turbine rotors schematically indicated at34. In another non-limiting example embodiment low pressure turbine 46includes about 3 turbine rotors. A ratio between number of fan blades 42and the number of low pressure turbine rotors is between about 3.3 andabout 8.6. The example low pressure turbine 46 provides the drivingpower to rotate fan section 22 and therefore the relationship betweenthe number of turbine rotors 34 in low pressure turbine 46 and number ofblades 42 in fan section 22 disclose an example gas turbine engine 20with increased power transfer efficiency.

FIG. 2 is a schematic perspective view of one embodiment of themid-turbine frame 57. The schematic view shown in FIG. 2 is a high levelconceptual view and is intended to illustrate relative positioning ofvarious components, but not the actual shape of various components. Themid-turbine frame 57 includes an outer frame case 70, an inner framecase 72, and a plurality of hollow spokes 74. In the example shown inFIG. 2, six hollow spokes 74 are distributed around the circumference ofthe inner frame case 72 to provide structural support between the innerframe case 72 and the outer frame case 70. In alternative embodiments,the mid-turbine frame 57 can have more or less than 6 hollow spokes. Theinner frame case 72 supports the rotor assembly via the bearing systems38 (shown in FIG. 1), and distributes the force from the inner framecase 72 to the outer frame case 70 via the plurality of hollow spokes 74as known.

FIG. 3A shows a portion of the mid-turbine frame 57 through whichcooling air is directed into a rotor cavity 60 via a swirler tube 62.The direction of entry of the cooling flow in the turbine is critical toensure that the cooling flow is directed at the correct location. Thesubject invention provides a mistake-proof mounting configuration toensure that the tube 62 is properly installed to direct cooling air asneeded.

Any type of swirler tube 62 or other cooling structures can be used withthe subject mounting configuration. One example of a swirler tube 62 isdescribed in co-pending application having Ser. No. ##/###,###, which isassigned to the same assignee as the subject invention and which ishereby incorporated by reference. In this configuration, the swirlertube 62 directs the cooling airflow into the rotor cavity 60 in thedirection of rotation of the low rotor to generate pre-swirling. Bypre-swirling the cooling air flow prior to entering the rotor cavity 60,the heat-up of the cooling air flow is reduced which lowers atemperature of the rotor cavity 60.

As shown in FIG. 3A, the mid-turbine frame 57 includes a plate portion64 having a forward face 66 and an aft face 68. The plate portion 64 ofthe mid-turbine frame 57 is attached to an engine inner case structure78

The engine inner case structure 78 is secured to the forward face 66 ofthe plate portion 64. The plate portion 64 includes at least one opening80 that is located radially outwardly relative to the inner case 78. Theopening 80 extends through the entire thickness of the plate portion 64to form a portion of the cooling flow path into the rotor cavity 60. Acooling structure, such as the swirler tube 62 for example, isassociated with the opening 80 and is configured to direct cooling flowin a desired direction into the cooling cavity, e.g. rotor cavity 60.

A bracket assembly 82 supports the swirler tube 62 and includes anattachment interface 84 to fix the bracket assembly 82 to themid-turbine frame 57. In the example shown, the bracket assembly 82comprises a forward bracket 86 positioned on the forward face 66 and anaft bracket 88 positioned on the aft face 68 of the plate portion 64.The swirler tube 62 is associated with the aft bracket 88.

The forward bracket 86 comprises a mounting flange portion 90 thatincludes the attachment interface 84 and a center boss portion 92 thatdefines a portion of the cooling flow path into the opening 80 in theplate portion 64. The center boss portion 92 defines a center axis andis associated with an inlet tube 94 (FIG. 3B) that receives cooling airflow from the compressor section 24. A tab portion 96 extends outwardlyfrom a peripheral edge of the mounting flange portion 90.

The aft bracket 88 comprises a mounting flange portion 98 that includesthe attachment interface 84 and a center boss portion 100 that receivescooling flow from the opening 80 in the plate portion 64. The centerboss portion 100 (FIG. 4) defines an opening and is associated with theswirler tube 62 that receives cooling air flow from the inlet tube 94. Atab portion 102 extends outwardly from a peripheral edge of the mountingflange portion 98.

The swirler tube 62 can be separately attached to the center bossportion 100, or the swirler tube 62 can be integrally formed with theaft bracket 88. The swirler tube 62 comprises a curved tube body 62 aextends to a distal tip 62 b from which cooling flow exits the tube 62.The cooling flow initially passes through the plate portion 64 in afirst direction that is generally parallel with axial flow through theengine. The curved tube body 62 a receives this flow and redirects theflow from this first direction defined by the inlet tube 94 to a seconddirection that is transverse to the first direction. In one example, theswirler tube 62 directs the cooling flow in a direction that isnon-parallel with the central axis defined by the inlet tube 94 toinitiate a swirling flow pattern about the central axis of the engine.

The plate portion 64 includes at least one first rib 104 extendingoutwardly from the forward face 66 and at least one second rib 106 (FIG.4) extending outwardly from the aft face 68. The ribs 104, 106 arelocated radially outwardly of the inner case structure 78. The ribs 104,106 comprise orientation features that cooperate with the tab portions96, 102 of the forward 86 and aft 88 brackets to ensure that the swirlertube 62 is only installed in one orientation relative to the mid-turbineframe 57.

The tab portions 96, 102 comprise a first orientation feature that isassociated with the bracket assembly 82 and the ribs 104, 106 comprise asecond orientation feature that is associated with the gas turbineengine structure, e.g. the mid-turbine frame 57. These first and secondorientation features cooperates with each other to prevent mis-assemblyof the swirler tube 62 in the gas turbine engine.

FIGS. 2-4 show one example of first and second orientations features;however, other types of features could also be used. For example, FIGS.5-6 show different example configurations where one of the first andsecond orientation features comprises a male feature and the other ofthe first and second orientation features comprise a female feature thatis received within the male feature to ensure proper installation.

FIG. 5 shows an example where the plate portion 64′ includes an opening,slot, or recess 120 that comprises a female feature. A forward bracket86′ includes a pin or hook portion 122 that extends outwardly from thebracket to comprise a male feature. The pin or hook portion 122 isreceived within the recess 120 when the forward bracket 86′ is in thecorrect orientation. The aft bracket 88 could be similarly configured toinclude these male and female features.

FIG. 6 shows an example where the plate portion 64″ includes an opening,slot, or recess 130 formed in the first rib 104″ that comprises a femalefeature. A forward bracket 86″ includes a tab portion 132 that extendsoutwardly from the bracket to comprise a male feature. The tab portion132 is received within the recess 130 when the forward bracket 86″ is inthe correct orientation. The aft bracket 88 could be similarlyconfigured to include these male and female features.

As discussed above, in the configuration shown in FIGS. 2-4, the ribs104, 106 cooperate with the tab portions 96, 102 to ensure that theswirler tube 62 is only installed in one orientation relative to themid-turbine frame 57. The attachment interface 84 of the forward bracket86 is located on the mounting flange portion 90 and comprises at leastone opening 140 configured to receive a fastener 142. In this example,there are a pair of openings 140 (only one can be seen) and anassociated pair of fasteners 142, wherein the openings 140 are locatedon opposite sides of the center boss portion 92.

The attachment interface 84 on the aft bracket 88 is located on themounting flange portion 98 and comprises at least one opening 144. Inthe example shown, there are a pair of openings 144 that are alignedwith the openings 140 in the forward bracket 86. The plate portion 64,which is sandwiched between the forward 86 and aft 88 brackets, alsoincludes a pair of openings 146 that are aligned with the openings 140,144. The fasteners 142 comprise a single set of fasteners that are usedto secure all three components together.

When properly installed, the mounting flange portions 90, 98 aregenerally flush with the respective forward 66 and aft 68 faces (FIG.2). FIG. 3B shows an incorrect orientation of the forward bracket 86.The bracket 86 has been rotated 180 degrees from the correct orientationposition. The tab portion 96 abuts against, or interferes with, thefirst rib 104 in this position which prevents the bracket 86 fromsitting flush with the plate portion 64. As such, the installer caneasily recognize that the bracket 86 is in the incorrect portion and canadjust the position accordingly.

FIG. 4 shows an incorrect orientation of the aft bracket 88. The bracket88 has been rotated 180 degrees from the correct orientation position.The tab portion 102 abuts against, or interferes with, the second rib106 in this position which prevents the bracket 88 from sitting flushwith the plate portion 64. As such, the installer can easily recognizethat the bracket 88 is in the incorrect portion and can adjust theposition accordingly.

As such, the subject invention provides a mistake proof mountingconfiguration for a cooling structure such as a swirler tube, forexample. This ensures that the entry of cooling air flow is in thecorrect direction, which is critical for providing the most effectivecooling. If the flow enters in an incorrect direction, the temperatureof the cooling air would be increased due to mixing and windage losses.If cooling air temperatures are increased due to windage, it can resultin lower life for the rotor or a reduced burst margin. The subjectinvention thus avoids these problems.

Further, should be understood that while the bucket assembly 82 is shownas mounting a swirler tube to a mid-turbine frame 57, the bracketassembly 82 could be used with other gas turbine engine structures. Forexample, the bracket assembly 82 could be used to mount other types ofcooling structures to direct cooling flow into any type of coolingcavity located throughout the gas turbine engine.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claims should bestudied to determine the true scope and content of this disclosure.

What is claimed is:
 1. A cooling structure for a gas turbine enginecomprising: a gas turbine engine structure defining a cooling cavity; acooling component configured to direct cooling flow in a desireddirection into the cooling cavity; a bracket supporting the coolingcomponent and having an attachment interface to fix the bracket to thegas turbine engine structure; a first orientation feature associatedwith the bracket; and a second orientation feature associated with thegas turbine engine structure, the first and second orientation featurescooperating with each other to ensure that the cooling component is onlyinstalled in one orientation relative to the gas turbine enginestructure.
 2. The structure of claim 1, wherein one of the first andsecond orientation features comprises a male feature and the other ofthe first and second orientation features comprises a female featurethat is received within the male feature.
 3. The structure of claim 2,wherein the male feature comprises a tab and the female featurecomprises a slot.
 4. The structure of claim 2, wherein the male featurecomprises a pin and the female feature comprises an opening.
 5. Thestructure of claim 1, wherein the first orientation feature comprises atleast one tab and the second orientation feature comprises a rib, andwherein the rib interferes with the tab to prevent mounting the bracketin an incorrect orientation relative to the gas turbine enginestructure.
 6. The structure of claim 1, wherein the bracket comprises aforward bracket positioned on one side of the gas turbine enginestructure and an aft bracket positioned on an opposite side of the gasturbine engine structure, and wherein the cooling component isassociated with the aft bracket, and wherein the forward and aftbrackets each include the first orientation feature.
 7. The structure ofclaim 6, wherein the gas turbine engine structure includes an opening todefine a portion of a cooling flow path, and further including an inlettube associated with the forward bracket to direct cooling flow throughthe opening and into the cooling component.
 8. The structure of claim 7,wherein the cooling component comprises a swirler tube that redirectscooling flow from a first direction defined by the inlet tube to asecond direction that is transverse to the first direction.
 9. Thestructure of claim 6, wherein the attachment interface includes at leastone fastener that secures both the forward and aft brackets to the gasturbine engine structure.
 10. The structure of claim 1, wherein the gasturbine engine structure comprises a mid-turbine frame structure. 11.The structure of claim 1, wherein the cooling component comprises aswirler tube.
 12. A gas turbine engine comprising: a mid-turbine framelocated axially between a first turbine and a second turbine, themid-turbine frame including an opening to define a portion of a coolingflow path into a rotor cavity; a cooling component configured to directcooling flow in a desired direction into the rotor cavity; a bracketassembly supporting the cooling component and having an attachmentinterface to fix the bracket assembly to the mid-turbine frame; a firstorientation feature associated with the bracket assembly; and a secondorientation feature associated with the mid-turbine frame, the first andsecond orientation features cooperating with each other to ensure thatthe cooling component is only installed in one orientation relative tothe mid-turbine frame.
 13. The gas turbine engine of claim 12, whereinthe opening in the mid-turbine frame defines a center axis, and whereinthe cooling component comprises a swirler tube that directs cooling flowin a direction that is non-parallel with the center axis.
 14. The gasturbine engine of claim 13, wherein the bracket assembly comprises aforward bracket positioned at the opening on one side of the mid-turbineframe and an aft bracket positioned at the opening on an opposite sideof the mid-turbine frame, and wherein the swirler tube is associatedwith the aft bracket.
 15. The gas turbine engine of claim 14, whereinthe first orientation feature comprises at least one tab and the secondorientation feature comprises a rib, and wherein the rib interferes withthe tab to prevent mounting the bracket in an incorrect orientationrelative to the mid-turbine frame.
 16. The gas turbine engine of claim14, wherein the first orientation feature is formed on both the forwardand aft brackets, and wherein the second orientation feature is formedon each side of the mid-turbine frame such that each of the forward andaft brackets can only be mounted on the mid-turbine frame in onemounting orientation.
 17. A method of installing a cooling structure ina gas turbine engine comprising: providing a gas turbine enginestructure that defines a cooling cavity and a cooling component that isconfigured to direct cooling flow in a desired direction into thecooling cavity; providing a bracket configured to support the coolingcomponent and to attach the bracket to the gas turbine engine structure;forming a first orientation feature on the bracket; forming a secondorientation feature on the gas turbine engine structure; and associatingthe first and second orientation features with each other to ensure thatthe cooling component is only installed in one orientation relative tothe gas turbine engine structure.
 18. The method of claim 17, whereinone of the first and second orientation features comprises a malefeature and the other of the first and second orientation featurescomprises a female feature, and including inserting the female featureinto the male feature to achieve the one orientation.
 19. The method ofclaim 17, wherein the one of the first and second orientation featuresinterferes with the other of the first and second orientation featuresto prevent improper installation.
 20. The method of claim 19, whereinthe gas turbine engine structure comprises a mid-turbine frame and thecooling component comprise a swirler tube.